Modern Advances in Fan and Compressor Design Using CFD

Riccardo Balin1 University of Colorado Boulder, Boulder, CO 80309

In this paper a literature review of the state of the art of fan and compressor blade design is presented. In particular, the role played by CFD simulations in the understanding of fully three-dimensional flow is reviewed, as well as the 3-D design features that can be studied with high fidelity simulations. The paper focuses on the effects of leading edge recambering, sweep, dihedral, and endwall contouring on the secondary flow structures typical of turbomachinery, in order to provide a guideline to the most efficient design solutions. The effect of blade sweep is also considered in the context of fan and compressor surge.

I. Introduction OMPRESSOR and fan design for air breathing machines applied to aircraft propulsion has made some drastic C leaps forward since the beginning of the jet era with the Jumo 004 and Whittle engines. At the core of the continuous development and research being carried out in this field is the everlasting search for high efficiency, low noise, durable and safe machines to improve air travel. In the world of commercial air travel, some of these aspects are extremely important, and can define the immense success or drastic failure of a design. Commercial air travel is a multi-trillion dollar business, driven mainly by airline companies and their desire to profit. In recent years, airline companies have been pushing aircraft and engine manufacturers to design more efficient, safer and quieter products. As a consequence, manufacturers are forced to stretch the design limits of their products and come up with new and creative designs to meet the requirements set by their customers. In addition, regulatory agencies such as the FAA and the European counterpart EASA are constantly raising the demands set on aircraft and engines with the goal of making travel safer and more comfortable. Some of these requirements involve the amount of production of nitrogen oxides during combustion, some involve noise levels at airports, and some involve the robustness of the engines to ingestion of foreign bodies. Focusing on the fan and compressor components of an air a) breathing engine, the requirements that drive most of the research in the field are minimizing losses, or equivalently maximizing efficiency, minimizing noise production, and increasing the reliability of the engine. Given the increase in fuel prices of the last couple decades, as well as the increased general awareness towards environmental issues, airlines are always demanding a reduction in the specific fuel consumption of the engines. Similarly, both airlines and regulatory agencies are demanding that the engines by quieter in order to make air travel and locations with dense air traffic more appeasable to the public. Engine reliability specific to fans and compressors can be thought as an increased surge and stall margin, thus reducing the possibility of engine failure. b) All of the aforementioned requirements (reduced fuel Figure 1. Comparison between a) modern consumption, reduced noise, increased stall margin) can be met 1 with blade and endwall design. Until recent years, blade design and b) legacy turbofan engine rotors.

1 Graduate student, Department of Aerospace Engineering Sciences, Student Member AIAA. 1

was very much limited to a two-dimensional analysis of the flow over an airfoil. The analysis would be iterated in order to achieve the airfoil profile which would produce the desired flow, and then the blade would be constructed by stacking the airfoils on top of each other. The two-dimensional profile of the blade could be optimized for different spanwise locations, but only a few locations were chosen for practicality. With this design methodology, three-dimensional effects such as sweep and dihedral could not be analyzed. Experimental tests could be conducted to study such 3-D effects, however this approach is expensive and time consuming. Regardless of these limitations that prevent full 3-D design of fan and compressor blades, the state of the art in the field has made incredible progress. Figure 1 shows a comparison between a modern high bypass ratio fan (Fig. 1a) and a legacy fan (Fig. 1b). The differences are obvious, and it is clear that blades are now fully three- dimensional. The main factor that allowed such progress is the advent of Computational Fluid Dynamics (CFD) and the development of turbulence models that can reproduce high fidelity results even for complex flows, while not requiring incredibly expensive simulations. More specifically, it is the combined effect of high speed computers with the development of accurate and efficient CFD codes that enabled fully 3-D flow analysis to be introduced into the design process, even at early stages. The fact that it is now possible to study a complex, fully 3-D flow in simulations with fast turn-around time means that blade design is no longer limited to airfoil profiling, but instead a number of other possibilities are introduced. Researchers have been exploiting this new technology to span the 3-D design space, and have been able to understand the 3-D properties of flows across fans and compressor, and how 3- D blades affect them. This paper consists of a literary review of the state of the art of 3-D fan and compressor blade and endwall design using CFD. It will focus both on characterizing the three-dimensionality of the flow across fans and compressors, and on the changes that can be made to the geometry in order to affect the flow properties. The rest of the paper is divided in to four sections, namely twist, sweep, dihedral and endwall contouring. The first three relate to changes to the blade geometry directly, while the last is focused more on the interaction between blade and endwall. In these sections, the effects of 3-D blade design on noise, losses and surge margin are outlined.

II. Twist One three-dimensional design feature that can be implemented in compressor and fan blades is geometric twist. Note that this feature is implemented in rotor blades, and not as often on stator blades. The main effect of blade twist is to ensure that every section of the blade operates at the desired incidence angle with the incoming flow. The relative wind seen by a blade airfoil is the vector sum of the incoming absolute velocity and the velocity of the blade generated by its rotation. The latter is a function of the radial distance of the airfoil section to the hub of the rotor, so the direction of the relative wind changes with radial distance. At the hub, where the blade tangential speed is small, the direction of the incoming Figure 2. Contours of relative velocity measured flow is almost axial, while at the tip it is almost experimentally at a rotor exit. 4 entirely in the tangential direction.2 In order to maintain the angle of attack of the blade section at good operating values, it is necessary for the blade to be twisted. An example of twist on a fan blade is seen in Fig. 1a. It is noticed that closer to the hub, the chord of the blade section is almost parallel to the axial direction in accordance with the direction of the relative velocity. Instead, the tip blade section is almost at 90° to the axial direction. One interesting difference between blade twist for fans and compressors is the fact that for a fan, the amount of twist is dictated by the design free stream Mach number, the design fan rpm and by the fan dimensions. However, a compressor blade is fairly short compared to a fan blade, especially at the last stages of a high pressure compressor. At these reduced dimensions, the height of the boundary layer that develops on the casing walls becomes significant relative to the blade length. This is particularly true because regions of secondary flow create at the endwall of a compressor due to tip vortices and possible separation of the flow on the blades of the previous stage. In addition, for rotating components, the boundary layer which forms on the surface of the blades is subject to significant centrifugal forces. This results in the boundary layer flow being pushed radially outward towards the trailing edge of the tip of the blades.3 This accumulation of low energy fluid at the tip of the blade promotes flow separation. Figure

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2 shows contours of velocity measured experimentally at the exit of a rotor stage, from which the secondary flow region at the junction between the tip of the blade and the casing endwall is clear. Because of this accumulation of low energy, secondary flow on the wall of the casing, the tip of rotor and stator blades see a reduction in axial velocity, which increases significantly the incidence angle. Consequently, it is desirable to twist the sections of compressor blades near the tip to achieve the desired incidence angle and loading on the whole blade.5 This type of blade twisting, which is specific to compressor blades and not fan blades, is given the name of leading edge recambering (LER). As the name suggests, LER involves more than a simple twist of the airfoil about its centroid. The airfoil camber, chord and thickness are redesigned to optimize performance in this region of the flow. It is Figure 3. Effect of LER on velocity contorus (original 6 extremely difficult to predict the physics of these blade on left, LER modified on right). secondary flow structures without the use of a CFD simulation. In fact, it can be stated that secondary flow effects on turbomachinery can only be understood with CFD analysis5, which is essential for proper LER design. Figure 3 shows the effects of LER optimal design obtained with CFD analysis. It can be seen that for the original blade (left picture), the boundary layer is retarded at the trailing edge, causing the wake extending from the blade to be relatively thick. This behavior is undesirable because large wakes cause flow distortion for the downstream stages. After LER, the boundary layer on the suction surface of the blade remains attached all the way to the trailing edge, causing a smaller wake. In addition to reducing flow distortion, LER was also able to reduce the stagnation pressure losses for the blade. Retarded boundary layers and wakes are both turbulent phenomena, therefore they generate a significant amount of entropy, which is seen in the flow as a drop in the total pressure.

III. Sweep Another 3-D geometric design feature is blade sweep. This concept was first applied to aircraft wings with the purpose of increasing the critical Mach number of the wing section and thus reduce drag at transonic and supersonic speeds. The pressure distribution around a cross section of a blade of infinite aspect ratio is determined solely by the component of the flow normal to the leading edge. For finite geometries, such as fan and compressor blades, this statement applies to a certain extent only. Nevertheless, the end result is that the velocity component normal to the leading edge is reduced by a factor approximately equal to the cosine of the sweep angle. This is precisely the effectiveness sweep, that is to reduce the effective flow velocity seen by the blade section. Both forward and aft sweep is possible, and accomplish the same aerodynamic effect. Sweep is dealt slightly differently for fan blades and compressor blades. In the case of fans such as the ones in high aspect ratio turbofans, both aft and forward swept blades produce the same effect of delaying shock formation on the suction surface. Nevertheless, other aerodynamic effects, as well as mechanical issues, must be taken into consideration when evaluating the benefits of aft and forward sweep. As stated earlier, the low energy fluid in boundary layers of rotating blades is forced radially out due to the centrifugal effects. This accumulation of low energy fluid at the tip of the blade promotes flow separation and causes additional losses and noise.7 A swept back blade has strong spanwise component of the flow velocity directed outwards towards the tip. This has the effect of increasing the accumulation of entrained fluid, therefore increasing the area of flow separation and losses. On the contrary, the spanwise flow on a swept forward blade is directed towards the hub, thus balancing the centrifugal force acting on the boundary layer and reducing the accumulation of low energy fluid. In addition, in a forward swept blade, the leading edge of the tip is located far in front of the mid-chord and away from the boundary layer drift motion. This concentrates the separated flow closer to the trailing edge of the tip, which is advantageous for two reasons. First, the size of the region of separated flow is smaller, and thus the stagnation pressure losses are reduced making the fan more efficient. Second, the loading of the tip of the fan is larger because flow is attached on a greater section of the blade.

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Another aspect to consider is the stability of the fan. The shocks generated on the suction surface of one blade propagate through the region of flow between the blades and end on the pressure surface of the adjacent blade. These passage shocks are a great source of losses, however it is not possible to remove them entirely. Instability in the fan occurs when the passage shock is exposed forward of the leading edge of the adjacent blade and sudden changes in the direction of the flow through the rotor occur, possibly resulting in the failure of the engine. Research performed by Rolls Royce on fan blades indicated that the line connecting the suction peaks on the suction surface of the blade, and thus the line connecting the locations at which the shock is formed, has a smaller sweep angle than that of the leading edge.8 Consequently, at the tip, the shock forms close to the leading edge for an aft swept blade, but it forms close the trailing edge for a forward swept blade, yielding a much safer design with a bigger margin for instability. Figure 4 illustrates this point.

a) b) Figure 4. Obloque shock location on a fan blade with a) sweep back and b) sweep forward.8 From the above discussion it is apparent that a forward swept blade provides significant advantages in reducing the compressibility losses at transonic speeds, and at the same time provides greater stability. Nevertheless, an entirely forward swept blade is not mechanically feasible in terms of the stresses sustained during operation without increasing the thickness of the blade significantly.8 This is because if the centroids of the blade sections are not radially aligned, bending moments act at the root of the blade, significantly increasing the stress. Therefore, a backward swept blade must be employed, but with a forward swept tip to increase stability and reduce accumulation of retarded fluid. Note that to align the centroids of the blade sections as much as possible, and hence reduce stresses, the first segment of the blade is forward swept too. This alternation between fore and aft sweep is also seen in the turbofan in Fig. 1a. In compressor blades, sweep is not only used for the reasons outlined previously for a fan, but also to control loading of the blade sections near the tip and the hub. Figure 5 illustrates how sweep affects the blade loading. Since there is no acceleration of the fluid at the endwalls in the direction perpendicular to the wall, the pressure gradient in the direction normal to the wall is close to zero. Near the leading edge of the blade close to the hub, the loading rapidly falls to zero moving away from the hub endwall along the blade surface because the blade ends (see lower left arrow in Fig. 5). Since the pressure gradient in the vertical direction is small, the loading on the leading edge of the blade close to the hub must also be small. Instead, at the trailing edge of the blade in proximity of Figure 5. Effect of sweep on blade loading near the hub, the loading tends to be increased because, due to endwalls.5 the small vertical pressure gradients, it takes the values of the more loaded region above it. Similar arguments are made for the blade section near the casing endwall. The effect of blade sweep of the type in Fig. 5 has thus the overall effect of shifting blade loading away from the leading edge for sections close to the hub, and towards the leading edge for sections close to the casing. The magnitude of this effect is seen in Fig. 6, where CFD data of the surface pressure distribution for a blade with the same sweep at the hub and tip as in Fig. 5, but no sweep in the middle section of the blade, is shown. Note that the data plotted in the figure is for a turbine blade, but the effects on compressors blades are the same.5 It is clear that relative to the distribution at the mid-span, which has no sweep, the distribution at the hub as a loading shifted towards the trailing edge, while the opposite is true for the distribution at the tip. Therefore, the ideal implementation of sweep in a compressor blade is to have forward sweep on both ends of the blade. This is to move the loading away from the

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leading edge in those sections, which can help to reduce the losses due to shock creation on the suction surface, but also to make the blade more tolerant to changes in the inlet angle. Because of the secondary flow regions that form on both the hub and casing endwalls, the inlet angle to the blade can vary frequently in these regions, meaning that a more tolerant blade can allow larger changes in angle without much effect on the loading. Recall that forward sweep at the tip also opposes the outward radial drift of secondary flow due to centrifugal motion, thus energizing the low-energy secondary flow at the wall and suppressing corner separation and losses.

IV. Dihedral Dihedral applied to compressor and fan blades can be defined as the shifting of blade sections in the direction perpendicular to the chord. Literature regarding dihedral design on fan blades was not found. A reason for this might be the fact that fans, because of their dimension, are much less affected by endwall secondary flows than compressor blades, where dihedral has the purpose of influencing endwall flow. One interesting aspect of blade dihedral is that its effects are not Figure 6. Static pressure distribution on a confined to a blade row, but instead it induces structures that blade with sweep as is Fig. 5.5 decay downstream of the blade row in a distance of the order of the blade height, thus affecting adjacent rows as well.5 One of the applications of dihedral is to change the spanwise variation of stage reaction. The stage reaction R is defined as the ratio of the enthalpy change across the rotor to the change in stagnation enthalpy across the entire stage. In the cases where the hub-to-tip ratio of the stage is very small, such as the last stages of a compressor, the root reaction can become very small because of the large flow velocities at the hub of the stator. 5 The reaction can be increased by leaning the rotor blades so that the pressure side is facing the hub. The blades impose a body force on the fluid as it passes through them which makes the streamlines curve, and when blades are aligned in the radial direction, this force has a component only in the tangential direction. When a dihedral angle is introduced, the force has a component in the radial direction as well and stream curvature is no longer in the tangential direction only. If the dihedral angle is such that the pressure surface of the blade is facing the hub, then the body Figure 7. Pressure distribution in a leaned force acts to increase the pressure on the hub surface between 5 blade with no spanwise variation in flow. blades. A diagram illustrating this effect for an ideal situation with no spanwise variation in flow is shown in Fig. 7. The increase in pressure at the hub, however, is not limited to the blade to blade section of the hub, and extends upstream and downstream of the blade row as well. This means that dihedral can be used to influence the hub pressure between stages. The increased pressure at the hub reduces the flow speed in that region (fluid velocity is inversely proportional to pressure), so the blade reaction at the hub is increased for downstream stages. For low aspect ratio blades, such as the ones in compressors, changes in dihedral angle can be considered as moving the blade inside an almost frozen pressure field. 5 Regardless of the dihedral, constant pressure lines between blades are very closely radial, with the pressure increasing moving from the suction surface of one blade to the pressure surface of the adjacent blade. These pressure fields are shown for three types of blades in Fig. 8. In the case of the straight lean, the root is moved to a region of higher pressure and lower velocity, while the tip is moved to a region of low pressure and high velocity. Consequently, the loading at the root is reduced, however at the tip there is a great increase in loading, which causes much stronger shocks to form on the suction side of the blade. This effect is undesirable in terms of efficiency because total pressure losses are larger for stronger shocks. The bade design at the bottom of Fig. 8, called a bowed design, moves both the root and the tip of the blade to a region of higher pressure and lower velocity. The resultant effect is that loading is reduced on both ends of the blade, and correspondingly it is increased at the mid-section. Bowed blades have a net effect of reducing endwall losses, however the precise mechanism by which this occurs is not known. One theory is that the shifting of the loading distribution to the mid-span region of the blade means that a larger fraction of the total work is being performed by

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the most efficient part of the blade. A second explanation is that bowed lean produces spanwise pressure gradients directed PS SS towards the mid-span of the blade. These pressure gradients would then help reduce the accumulation of low energy fluid at the endwalls, and thus reduce local separation near the endwalls. One additional advantage of bowed dihedral is to reduce the tip losses. This is because tip losses are caused by an imbalance of the pressure between the suction side and the pressure side of the blade, and is proportional to the magnitude of that difference. In bowed blades, the loading at the tip is reduced, and consequently the difference in pressure between the two sides of the blades is also reduced, resulting in smaller tip losses. 5

V. Endwall Contouring Endwall contouring refers to changing the profile of the hub and casing endwalls between blade rows so that it is not flat and axisymmetric, but it becomes profiled with slight “valleys and hills.” The main purpose of this design strategy is to affect the corner losses that generate at the junctions between blades and endwall. In Ref. 9, profiled endwall Figure 8. CFD computed static pressure (PEW) design is used to reduce hub corner stall on contours between blades with different 5 compressor stators. The process of designing the proper dihedral. contours is done combining a CFD simulation into optimization codes, where the code itself goes through the process of designing an endwall profile, running a CFD simulation on the current geometry, and using a series of

Figure 9. CFD computed particle paths and reverse flow isosurfaces for different blade sedigns.9 specified indicators to improve the profile and repeat the process. In Ref. 9, the authors implement this optimization process and develop a series of endwall profiles, some of which are more successful than others. The geometries are then compared to a regular 2-D blade design and a 3-D blade design, and some of the results are shown here in Fig. 9 and Fig. 10. Figure 9 shows how PEW on the wall of the hub affects the flow in the corner region. From the flow around the simple 2-D designed blade, it is seen that there is a large region of separated flow, which is caused by the interaction of the secondary flows that exist at the endwalls and the blade. It is clear that both the 3-D blade design and the PEW design significantly reduce the size of this region. The consequences in terms of losses are significant, and are shown in Fig. 10, along with an explanation of the process by which PEW and 3-D blades can reduce separation. In the top image, the horseshoe vortex on the suction side of the blade that is created by the leading edge is shown and labeled as A. It can be seen that the location and size of the vortex is similar for all three blade designs. At 30% of the chord, however, difference in the designs is seen. The for the 2-D blade, the horseshoe vortex, which is a region of low-energy, low-velocity fluid, grows in size but its location does not change. On the other hand, for the 3-D and PEW design, the vortex is pushed towards the blade. In addition, it does not grow in size as much as for the 2-D 6

blade, especially in the case of the 3-D blade. The same tends are seen at 40% of the chord, where the 3-D and PEW designs have almost suppressed the vortex and pushed it onto the blade. Instead, the 2-D design is showing a region of separated flow. The end effect is seen at 70% of the chord, where the 2-D design shows a massive region of separation and loss, while the other two designs do not. It is interesting to point out that the 3-D design performs better than the PEW design. Nevertheless, the improvements generated by PEW are clear, and if PEW is combined with 3-D blade design, the effects can be greater than those of 3-D design alone. It is thus clear that in order to reduce corner separation, and therefore corner losses, the region of low-energy fluid created by the shedding of the horseshoe vortices around the leading edge and residing on the endwalls of the casing or the hub must be sucked towards the blade. This behavior is achieved by creating spanwise pressure gradients that move low-energy fluid towards the blade, and then further towards the blade mid-span, where they can be re-energized. Because of the large number of blades in an aircraft engine, small reductions in corner losses can result in a significant increase in efficiency of the engine as a whole.

Figure 10. CFD computed contours of total pressure loss coefficient at several axial planes for different blade designs.9 To show the overall effect of PEW design on a multistage compressor, a CFD simulation on a six stage high pressure compressor was performed by the authors of Ref. 9. The results were then compared between the compressor having 3-D designed blades and 2-D designed blades plus PEW. The authors claim that the same design point compressor efficiency was achieved between the two cases, thus indicating the competitiveness of PEW design. However, PEW design has one main disadvantage; the fact that endwall profiles are of the order of 0.7 mm, meaning that manufacturing the endwall to the required accuracy is extremely difficult.

VI. Conclusion This paper reviewed some of the three-dimensional design features that can be implemented to fan and compressor blades in order to reduce losses, reduce noise and increase the safety of the engine. All of these 3-D 7

features and all of the insight that is now available on secondary flows in turbomachinery were brought about by the advent of CFD codes and turbulence models. This is still a relatively young field, which is limited by factors such as manufacturing and materials. Some of the designs that are obtained from CFD simulations and geometry optimization are currently not feasible or very expensive and therefore not worth the investment for manufacturers. An example of this is PEW design. In the future, however, new technologies in other fields will allow engineers to exploit these 3-D design features even more to keep increasing efficiency of fans and compressors. Moreover, the increase in computer power of then next years will allow larger and more complex simulations to be carried out, thus improving the design and understanding of turbomachinery.

References 1Kantha, L., “Lecture 12,” ASEN 4012: Foundations of Propulsion, University of Colorado Boudler, Fall 2014. 2Anderson, J. D., Introduction to Flight, McGraw-Hill, New York, 1989, Chaps 5, 9. 3Mohammed, K. P., Prithvi Raj, D., “Investigations on Axial Flow Fan Impellers With Forward Swept Blades,” Journal of Fluids Engineering, paper No. 77-FE-1, Semptember 1977, pages 543-547. 4Gallimore, S. J., et al, “The Use of Sweep and Dihedral in Multistage Axial Flow Compressor Blading – Part I: University Research and Methods Development”, Journal of Turbomachinery, Vol. 124, Oct. 2002, p. 521-532. 5Denton, J. D., Xu, L., “The exploitation of three-dimensional flow in turbomachinery design”, Proceedings of the Institution of Mechanical Engineers, Vol. 213 Part C, 1999, p. 125-137. 6Shahpanar, S., Caloni, S., “Aerodynamic Optimization of High-Pressure Turbines for Lean-Burn Combustion System”, Journal of Engineering for Gas Turbines and Power, Vol. 135, May 2013. 7Agboola, F. A., Wright, T., “The Effects of Axial Fan Noise Control by Blade Sweep on the Radial Component of Velocity,” AIAA, A99-27841, 1998. 8Rowlands, P. A., Rolls-Royce PLC, London United Kingdom, U.S. Patent No. 6,071,077, filed 6 Jun. 2000. 9Harvey, N. W., Offord, T. P., “Some Effects of Non-Axisymmetric End Wall Profiling on Axial Flow Compressor Aerodynamics. Part II: Multi-Stage HPC CFD Study”, Proceesings of ASME Turbo Expo 2008: Power for Land, Sea and Air, Jun. 2008.

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